JP2742998B2 - Cooling air supply manifold - Google Patents

Description

Description: TECHNICAL FIELD The present invention relates to a structure for supplying cooling air to a turbine rotor of a gas turbine engine. BACKGROUND OF THE INVENTION Gas turbine engine rotors are provided with air supplied to a radially inner portion of the rotor by a manifold structure that discharges cooling air having a tangential velocity component selected to match the angular velocity of the rotor. Often, it is cooled by the flow of water. For example, 1984 3
The structure described in U.S. Pat. No. 4,435,123, issued on May 6, is mounted in a gas turbine engine to support an annular sealing surface or the like that seals between various parts of the engine. Often. The manifold structure receives cooling air from a pressurized annular stream supplied from an upstream compressor section. As can be appreciated by those skilled in the art, uniformity of the discharged cooling air achieves favorable cooling effects on the radially outer structures and blades of the turbine rotor exposed to the heated combustion products. Is a major factor in In addition to airflow uniformity, changing the operation of the cooling system or detecting plugging or other flow anomalies requires sensing the pressure in the cooling airflow space adjacent to the turbine rotor. is there. Also, those familiar with the development of gas turbine engines will recognize that the first stage cooling requirements of a turbine in a particular engine may be reduced during the life of the engine as the engine structure is upgraded to increase or decrease power output. It will be understood that it often changes. In the conventional cooling air manifold structure, it is necessary to change the size of the air flow path and its holes in order to accommodate the change in the cooling demand of the turbine rotor. A plurality of similar incompatible parts are required. Similarly, due to changes in rotor cooling requirements for a particular engine at the site, such as changes in engine construction materials and increased service life, the manifold currently used for that engine can be removed to provide the desired cooling airflow. It will be necessary to replace other manifolds that are manufactured to supply. Accordingly, what is needed in the art is a method that can provide a uniform air flow to a turbine rotor and is easily adapted to provide cooling air at various flow rates without replacement. It is a manifold to get. SUMMARY OF THE INVENTION It is therefore an object of the present invention to provide an airflow manifold that supplies an annular swirling flow of cooling air to a region adjacent a radially inward surface of a rotating turbine rotor disk such as a gas turbine engine. is there. It is another object of the present invention to provide an airflow manifold configured to distribute cooling air uniformly to an area adjacent a face of a turbine rotor disk. Yet another object of the present invention is to provide a means for mounting a manifold in a gas turbine engine that does not disrupt the flow of air through the manifold. It is yet another object of the present invention to provide a means for adjusting the flow rate of air through a manifold in response to turbine rotor disk cooling requirements. Yet another object of the present invention is to provide an integral pressure tap for detecting the outlet pressure of the manifold and the flow rate of cooling air. According to the present invention, a cooling air manifold that tangentially injects a swirling flow of cooling air into a radially inward portion of a rotating turbine rotor disk has an annular, pressurized, substantially axial flow. It consists of a plurality of identical channels adjacent to each other that receives cooling air from the air flow and guides the air radially inward to the nozzles of the manifold, and the cooling air guided to the nozzles is tangential to the rotating disk. Released. The manifold according to the invention is substantially symmetrical about the axis of rotation of the rotor, the flow path being defined by two substantially truncated cone-shaped walls having a plurality of flow dividers between them. Has been determined. Mounting of the manifold is performed by increasing the thickness of the flow divider to receive mounting bolts, so that the manifold can be secured to the engine case, i.e., the engine frame, without disturbing the airflow flowing through it. . The flow divider is curved in a region adjacent the outlet of the manifold to form a discharge nozzle that accelerates the discharged cooling air and provides a tangential velocity component to the cooling air. Each flow path does not have a conventional common air inlet and plenum structure that can create an internal pressure loss of fluid and unbalanced air flow. Another feature of the manifold according to the present invention is that the flow rate of air passing through the manifold can be adjusted without changing the configuration of the entire manifold. Such adjustment of the flow rate is achieved in the present invention by providing a flat surface for receiving the corresponding flow blocking plate at a location adjacent to the inlet of each flow path. The blocking plate blocks the flow of air entering the manifold, thereby providing an easy means to modify the cooling performance of the air flow. If additional airflow is required, a truncated cone-shaped wall is provided with an increased thickness area near the turbine rotor, through which a flow control hole is provided. Perforations are made, whereby a portion of the cooling air is discharged from the manifold, bypassing the discharge nozzle. The manifold also provides a secondary flow of cooling air between an axially flowing pressurized air flow and a radially inward portion of the turbine blade mounted on the rotor disk. A plurality of beveled holes are formed in a radially outward peripheral flange formed in a third truncated conical wall adjacent the turbine inlet. These beveled holes emit a secondary air flow tangential to the disks and blades to prevent hot combustion gases from flowing radially inward beyond the plane of the disk. Yet another feature of the manifold according to the present invention is that
A pressure tap passage extending between an engine space for receiving air discharged from a nozzle of the manifold and an upstream pressure port for connecting to a pressure detecting means is provided. Therefore, the cooling air discharge pressure is detected at a position close to the rotating disk without disturbing the cooling air flow in the manifold. Hereinafter, the present invention will be described in detail with reference to the accompanying drawings with reference to the accompanying drawings. FIG. 1 is a sectional view showing a portion near a turbine rotor stage in a gas turbine engine according to an embodiment. The turbine rotor disk 10 and the blades are cooled by a flow of cooling air 14 flowing radially outward between the annular side plate 16 and the turbine rotor disk 10.
The flow of cooling air 14 is discharged from the nozzle 18 of the annular cooling air manifold 20 according to the present invention. The cooling air manifold 20 is adapted to receive cooling air from an annular, substantially axially flowing stream of pressurized cooling air 22 flowing radially inside the inner burner liner 24. The cooling air 22 is a radially extending dirt reflector 26
And flows into a plurality of channels 28 formed in the manifold 20. The inlet 30 of each channel is surrounded by a flat surface 32 for receiving a flow blocking plate, as described below. Rotary seals 34 and 34 are provided between the manifold and the disc and between the manifold and the side plate, respectively.
36 are placed and these seals are
This prevents the cooling air 14 discharged from the space 60 adjacent to 11 into the low-pressure region of the engine from leaking. FIG. 2 shows a manifold 20 according to the invention removed from the engine so that features other than those described above can be clearly understood.
FIG. A blocking plate 38 is shown in a position that covers a portion of the inlet 30, thereby restricting the flow of air entering the flow passage 28. The manifold 20 comprises a substantially frusto-conical first wall 40 and a spaced-apart frusto-conical second wall 42, which are disposed therebetween. The individual flow paths 28 are formed in cooperation with the plurality of flow dividers 46. The first and second walls extend radially inward and axially downstream from the inlet 30 to the nozzle 18. A truncated conical third wall 44 is the entrance to the channel 28
A peripheral mounting flange 66 extends radially outward and downstream from the vicinity of 30 and supports the rear end of the burner liner 24. 3 and 4 best illustrate the flow of cooling air flowing through the flow path 28. Each flow path 28 is a flow divider 46
Are separated from the adjacent flow path. Unlike the structure of the conventional manifold, the manifold 20 according to the present invention
Does not mix or distribute the cooling air before the received cooling air is released from the nozzle 18. Each flow path 28 has an inlet 30 and a nozzle 18 respectively, thereby providing a completely defined flow path without any interruption for the passage of cooling air. The radially inner portion of each flow divider 46 is circumferentially inclined to form a plurality of tangentially directed nozzles 18 that provide the desired velocity and swirl to the discharged cooling air 14. doing. The manifold 20 according to the present invention is provided with each flow divider.
46 is secured to the engine frame 48 by a plurality of mounting bolts 50 extending axially through corresponding mounting holes 52 provided in the boss region 54 having an increased thickness (see FIG. 1). ). By providing each flow divider 46 with a boss region 54 having an increased thickness, the manifold 20 according to the present invention can be mounted on the engine frame 48 without disturbing or separating the flow of cooling air passing through each flow passage 28. Can be fixedly attached to Unlike the conventional structure, in which the flow of air received through a plurality of inlets is mixed in a plenum region in the manifold and then discharged through a plurality of nozzle holes, the manifold 20 according to the present invention provides A carefully configured and fully defined flow path is provided for directing each portion of the cooling airflow flowing therethrough. The uniform flow path thus enables a uniform supply of air, which could not be achieved with the conventional manifold structure. Due to the double wall and divider structure of the manifold 20,
Thin and lightweight walls can be used without reducing the structural strength of the manifold as compared to conventional plenum configurations. The extra thickened boss region 54 serves the dual functions of locally strengthening the manifold 20 and splitting the cooling air flow between adjacent flow paths 28, thus providing extra independent mounting. The need for a structure is avoided and a reduction in weight is achieved compared to conventional manifolds. As mentioned above, to accommodate the cooling demands of the turbine rotor at various power levels over the life of the associated gas turbine engine, the flow of cooling air through the manifold is varied collectively or locally. It is necessary. Such a change may be achieved by providing one or more blocking plates 38 covering a portion of the inlet 30 of the flow path, as best shown in FIG. The blocking plate may be secured by welding or other means well known in the art, and
It is sized to introduce an appropriate amount of air into it. A slight flow adjustment and a slight increase in the total flow may be provided by the flow control boss 56 shown in FIGS. The flow control boss 56 is a portion where the thickness of the frusto-conical first wall 40 of the manifold is increased, and the flow passage 28
Some of the cooling air in the bypass the corresponding nozzle 18,
Flow control holes 58 are drilled as needed to allow flow into the turbine disk space 60 adjacent to the rotary seal 36 between the side plate and the manifold. By properly setting the size of the flow control hole 58, the flow rate of the bypass cooling air passing through the hole matches the leakage amount of the cooling air that is considered to pass through the seal 36, thereby discharging the cooling air from the nozzle 18 of the manifold. May be controlled to improve the cooling effect of the cooling air 14 flowing in the radial direction. Additional cooling of the radially inner portion of the blade 12 is provided by a plurality of angled holes 62 provided in the radially outer periphery of the third wall 44. The slanted holes 62 shown in FIGS. 2 and 7 discharge secondary cooling air tangentially adjacent the upstream side of the turbine rotor disk 10 so that hot combustion gases can escape from the turbine blades. It is provided in such a direction as to prevent it from flowing inward in the radial direction through the platform 64 (see FIG. 1). This inclined hole 62 is a flange
It is drilled in 66 and has a groove 68 formed in the manifold to facilitate the drilling for accessing a working tool. According to the double wall structure of the manifold 20 according to the present invention,
A uniform flow of cooling air 14 can be formed in the area adjacent to the rotating disk 10, but with a simple pressure tap hole for detecting the pressure of the flow of cooling air 14 in the turbine disk space 60 and thus Can not do. In the manifold 20 according to the present invention, as shown in FIG. 5, an internal pressure tap passage 70 that fluidly connects the turbine disk space 60 and a pressure tap hole 72 provided on the upstream side surface of the manifold. , The necessary pressure detection function of the related art is maintained. Pressure tap passage 70
Are formed in the manifold and form a pair of flow dividers.
It is located in the middle of 46 in the circumferential direction. In the illustrated embodiment, the pressure tap holes 72 are shown as being at the same radial position as the mounting bolts 50, but at various locations on the upstream side of the manifold and pressure sensing means (not shown). ) May be provided at a position that can be appropriately connected to Thus, the manifold 20 according to the present invention is an integral, adjustable cooling air structure which is very suitable for providing a uniform flow of cooling air on the upstream side 11 of the turbine rotor disk 10 in a gas turbine engine. Although the present invention has been described in detail with reference to specific embodiments, the present invention is not limited to such embodiments, and various other embodiments are possible within the scope of the present invention. That will be apparent to those skilled in the art.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 shows a turbine rotor disk, a discharge end of a combustor,
FIG. 2 is a sectional view along an axis showing a manifold according to the present invention. FIG. 2 is a sectional view showing the manifold according to the present invention separated from the surrounding engine structure. FIG. 3 is a perspective view showing a part of the manifold according to the present invention. FIG. 4 is a perspective view showing adjacent flow paths of the manifold with the frusto-conical upstream wall removed. FIG. 5 is a sectional view of the manifold passing through the pressure tap passage. FIG. 6 is a partial sectional view showing an inlet of one flow path in a state where a blocking plate is attached. FIG. 7 is a partial cross-sectional view showing cooling holes provided in a radially outer peripheral portion of the manifold. FIG. 8 is a partial cross-sectional view showing a flow rate adjusting boss and a hole provided therein. 10… Turbine rotor disk, 12… Blade, 14 ……
Cooling air, 16 ... Side plate, 18 ... Nozzle, 20 ...
Manifold, 22 ... Cooling air, 24 ... Burner liner, 26
…… Dirt deflector, 28 …… Flow path, 30 …… Inlet, 32…
… Flat surface, 34, 36 …… Seal, 38 …… Blocking plate, 40…
… First wall, 42 …… second wall, 44 …… third wall, 46 …… flow divider, 48 …… engine frame, 50… mounting bolt, 52… mounting hole, 54… boss Area, 56: Flow control boss, 58: Flow control hole, 60: Turbine disk space, 62
…… inclined hole, 64 …… platform, 66 …… flange, 70 …… pressure tap passage, 72 …… pressure tap hole

Claims (1)

(57) [Claims] A cooling air supply manifold for supplying an annular swirling flow of cooling air to one side of the rotating turbine disk (10), radially outward and axially upstream from a position adjacent to the turbine disk (10); A substantially truncated cone-shaped first wall (40) extending therefrom and a substantially truncated cone-shaped spaced-apart radially inward and axially upstream of the first wall. A second wall (42), a plurality of cooling air flow paths (2 extending between the first wall and the second wall and between the two walls);
8) forming a plurality of flow dividers (46), each of the cooling air passages having an inlet (30) and an outlet, wherein the radially inner end of each of the flow dividers is A plurality of inclined cooling air discharge nozzles (18) are formed at the outlet of the cooling air flow path so as to be inclined in the circumferential direction with respect to the rotation axis of the turbine disk, thereby forming the cooling air flow path (28). Each of them forms a passage individually communicating from the cooling air inlet (30) belonging thereto to the cooling air discharge nozzle (18), and the upstream end of each of the flow dividers (46) has an increased thickness. A boss area (54), wherein the boss area has mounting bolts (50);
A cooling air supply manifold provided with a mounting hole (52) through which air passes.